Aircraft manufacturers continuously attempt to improve aircraft performance by reducing both weight and manufacturing costs while maintaining or improving structural strength. One well-known method for increasing aircraft performance is to reduce airframe weight through the use of state-of-the-art materials, such as composites, having relatively high strength-to-weight and stiffness-to-weight ratios. Composite materials are generally described as being materials that include reinforcing fibers, such as graphite fibers, embedded in a polymeric matrix, such as an epoxy resin. Such materials will hereinafter be referenced as “fiber-reinforced composite” materials.
A variety of composite designs are being proposed on an on-going basis by a number of aircraft manufacturers.
U.S. Pat. No. 4,310,132 discloses a built-up fuselage formed by skins in composite materials stiffened by integral stringers shaped with “J” section, crossed by metallic frames. The frames are cut in order to permit the passage of uninterrupted stringers through cut-outs. The stringers are attached to the frames using clips and mechanical fasteners.
U.S. Pat. No. 5,242,523 discloses a bonded fuselage, totally made of composite materials, with skins and stringers shaped with an “omega” cross-section, crossing the frames through cut-outs, wherein each stringer has pressure bridges provided by additional parts.
U.S. Pat. No. 7,134,629 discloses a built-up fuselage, totally made of composite materials, formed by integral skins stiffened by stringers shaped with an “omega” cross-section crossing the frames through cut-outs.
Considerable disadvantages exist however in the known prior art, which requires the installation of the stringers and frames. Manufacturing stringers with different geometries is very expensive. Another disadvantage is that some stringer installation by riveting is time-consuming and cost-demanding. Such a method creates a great number of holes through the outer skin, which always increases the potential for corrosion and fatigue problems. Another concern is the intersection where the stringer passes the frames. Specifically, according to the current state of the art, each of the frames can include a base portion and an upstanding portion projecting away from the base portion wherein the upstanding portion can include a plurality of openings (colloquially known as “mouse houses”) through which the continuous stringer passes. Several problems can rise from the cut-outs, such as stress concentration, leading to small cracks which may grow in a straight line and extend through the complete section, spoiling the strength of the frame.
As conventional methods for manufacturing airframes in composite materials often require expensive tooling and labor-intensive assembly procedures it would therefore be highly desirable to develop new panel structures with reduced weight and increased cost efficiency in the manufacturing process. It is towards fulfilling such needs that the present invention is directed.
The present invention is therefore directed generally toward structural panels for use in manufacturing aircraft and other structures. The present invention relates generally to a monolithic integrated structure including a skin, a stiffener and a frame without cut-outs and a method of manufacturing the said structure.
According to some embodiments, a monolithic structural panel formed entirely of a fiber-reinforced composite material will comprise a skin, and frame and stinger members attached to the skin, wherein intersections between the frame and stringer members include cross-plied laminae of fiber-reinforced composite material. In some embodiments, the frame and stringer members are substantially mutually orthogonal to one another. The intersections between respective frame and stringer members comprise preformed modular units, wherein the modular units are structurally united to one another by lengthwise flanges formed of laminae of fiber-reinforced composite material. Bridge elements (preferably generally V-shaped) may be provided at the intersections between the frame and stringer members, the bridge elements are joined to the stringer members by means of the lengthwise flanges.
A one-piece structural panel for aircraft according to other embodiments of the invention may comprise mutually orthogonal frame and stringer members intersecting at a monolithic junction and formed entirely from a fiber-reinforced composite material.
In other embodiments, methods of making a structural panel are provided which include joining frame and stringer members to a skin, wherein each of the skin and frame and stringer members is formed of a fiber-reinforced composite material, and wherein intersections between the frame and stringer members are formed by cross-plying laminae of fiber-reinforced composite material. In some preferred embodiments, the frame and stringer members are mutually orthogonal to one another.
In certain embodiments, methods are provided to fabricate an aircraft structural panel by providing a preform of mutually orthogonal frame and stringer members formed of a fiber-reinforced composite material and having a monolithic junction therebetween, positioning the preform on a skin formed of a fiber-reinforced composite material, and thereafter co-curing the preform and the skin to form an aircraft structural panel.
These and other aspects and advantages will become more apparent after careful consideration is given to the following detailed description of the preferred exemplary embodiments thereof.